It is known practice for selected gas turbine engine components, especially in the combustor turbine interface, to be internally air cooled by a supply of air bleed from a compressor off-take. Such cooling is necessary to maintain combustor component temperatures within the working range of the materials from which they are constructed.
With reference to EP 2 230 456 A2, especially to FIG. 1 and FIG. 2 of that document, a typical gas turbine includes a transition piece by which the hot combustion gases from an upstream combustor as represented by the combustor liner are passed to the first stage of a turbine represented at item 14 in EP 2 230 456 A1. Flow from the gas turbine compressor exits an axial diffuser and enters into a compressor discharge case. About 50% of the compressor discharge air passes through apertures formed along and about a transition piece impingement sleeve for flow in an annular region or annulus (or, second flow annulus) between the transition piece and the radially outer transition piece impingement sleeve. The remaining approximately 50% of the compressor discharge flow passes into flow sleeve holes of an upstream combustion liner cooling sleeve (not shown) and into an annulus between the cooling sleeve and the liner and eventually mixes with the air in annulus. This combined air eventually mixes with the gas turbine fuel in a combustion chamber.
With reference to EP 1 426 558 A2 a transition piece assembly includes a transition piece and a surrounding impingement flow sleeve attached to the turbine end of the liner.
The impingement flow sleeve that is formed with a plurality of impingement cooling apertures by which compression cooling air enters the plenum created between the impingement flow sleeve and the outer surface of the transition piece to interact with the dimples on the cold side surface thereof. Thus, the flow sleeve assures that the convection cooling air is directed as desired along the transition piece from the turbine end to the combustor liner end, over the dimpled surface, increasing the heat transfer coefficients and reducing the temperature of the transition piece.
Generally, prior art sequential liner cooling concepts feature a backside cooled wall incorporating impingement cooling. Other cooling schemes utilize a certain length where the sequential liner is cooled with convective cooling techniques, for example smooth, rough or turbulated walls. Usually, the first impingement rows without any cross-flow in the cooling channel are located towards the turbine interface.
The mass flow of the single impingements jets are defined by the local driving pressure drop. The cooling mass flow is fed from the compressor exit. The impingement cooled section is then followed by convective cooling techniques. As a result the pressure drop of such a system is always an addition of the single pressure drop contributors.
Generally, higher engine gas temperature have led to increased cooling bleed requirements resulting in reduced cycle efficiency and increased emission levels. To date, it has been possible to improve the design of cooling systems to minimize cooling flow at relative low cost. In future engine temperatures will increase to levels at which it is necessary to have complex cooling features to maintain low cooling flows.
Additionally, referring to known solutions to lower the sequential liner pressure drop is to use essentially effusion cooling techniques, in the way that the cooling system is parallel to the combustion system. Drawback is the increased leakage into the combustion system, which is associated with higher emissions.
U.S. Pat. No. 5,802,841 A discloses a transition piece having one end adapted for connection to a gas combustor and an opposite end adapted for connection to a first turbine stage. Said transition piece having at least one external liner and at least one internal liner and the internal liner forms the hot gas flow channel. The first section of the transition piece assembly upstream of a first turbine stage having a plurality of cooling apertures. Cooling medium through the cooling apertures enters the plenum which is created between external and internal liners and a cooling medium flows along at least the first section of the transition piece assembly. At least one second section upstream of the first section with respect to the hot gas flow having at least one additional air inlet system. The additional air inlet system of the second section (FC) is designed in the manner that the cooling medium is discharged into at least one air plenum created between external and internal liners in different direction with respect to the hot gas flow in two different directions, namely an upstream and downstream one.
More pertinent material is evident from the following documents:
EP 2 378 200 A2; EP 2 148 139 A2; EP 0 239 020 A2; US 2006/283189 A1; JP H09 41991 A; US 2014/109577 A1.